Aircraft design for air to air refueling

ABSTRACT

An aircraft is provided including an airframe, an extending tail, and a counter rotating, coaxial main rotor assembly including an upper rotor assembly and a lower rotor assembly. The upper rotor assembly and the lower rotor assembly each include a plurality of blades. A translational thrust system is positioned at the extending tail and provides translational thrust to the airframe. The plurality of blades of the main rotor assembly extend beyond a nose of the aircraft by about 13 inches.

CROSS-REFERENCE TO RELATED APPLICATION

This application claims the benefit of U.S. provisional patentapplication Ser. No. 62/058,424 filed Oct. 1, 2014, the entire contentsof which are incorporated herein by reference.

BACKGROUND

The subject matter disclosed herein relates generally to rotary wingaircraft and, more particularly, to a dual rotor, rotary wing aircraft.

Aerial refueling or “in-flight” refueling has long been used to increasethe endurance of an aircraft. In general terms, a receiving aircrafttypically follows a tanker aircraft and engages a fuel transfer devicedeployed from the tanker aircraft. Following the engagement, fuel istransferred from the tanker aircraft to the receiving aircraft throughthe transfer device. After the transfer of a desired amount of fuel, thereceiving aircraft disengages the fuel transfer device and departs.

In one commonly used aerial refueling system, referred to as the “probeand drogue” system, a retractable hose is deployed from the tanker thatincludes a drogue coupled to a trailing end of the hose. The drogueincludes an internal valve that interrupts a flow of fuel to thereceiving aircraft when the drogue is disengaged. The receiving aircraftincludes a probe that extends forwardly and is configured to engage thevalve positioned within the drogue so that the receiving aircraft mayreceive fuel from the tanker.

In-flight refueling of rotary wing aircrafts, such as helicopters, hasbeen limited to operations in which the forwardly projecting fuel probecan engage the drogue trailing behind a fixed wing tanker aircraftcapable of flying at a minimum control speed in the range of the maximumspeed of the helicopter. However, few fixed wing tanker aircrafts canmaintain a minimum control speed within the maximum speed range of mosthelicopters. To prevent interference between the tanker aircraft and thehelicopter, the fuel probe must extend a significant length from thenose of the helicopter, beyond the rotor blades of the main rotorsystem. As a result of this extended length, engagement of the fuelprobe and the drogue in-flight is difficult to achieve. In addition, asa result of the extended length of the probe, the rotor blades maycontact the probe when the aircraft is in certain orientations.

SUMMARY

According to an embodiment, an aircraft is provided including anairframe, an extending tail, and a counter rotating, coaxial main rotorassembly including an upper rotor assembly and a lower rotor assembly.The upper rotor assembly and the lower rotor assembly each include aplurality of blades. A translational thrust system is positioned at theextending tail and provides translational thrust to the airframe. Theplurality of blades of the main rotor assembly extend beyond a nose ofthe aircraft by about 13 inches.

In addition to one or more of the features described above, or as analternative, in further embodiments a refueling probe is mounted to theairframe and fluidly connected to a fuel tank of the aircraft. Therefueling probe id configured to form a connection with a complementarycomponent to form a fuel flow passage there through.

In addition to one or more of the features described above, or as analternative, in further embodiments a first end of the refueling probedoes not extend beyond the plurality of rotor blades.

In addition to one or more of the features described above, or as analternative, in further embodiments the refueling probe is configured tomove between a retracted position and an extended position. The movementbetween the retracted position and the extended position does notinclude telescoping.

In addition to one or more of the features described above, or as analternative, in further embodiments in the retracted position, therefueling probe is housed within a cavity formed in the airframe nearthe nose.

In addition to one or more of the features described above, or as analternative, in further embodiments in the extended position, therefueling probe has a reduced impact on drag of the aircraft.

In addition to one or more of the features described above, or as analternative, in further embodiments the aircraft is configured toconnect the refueling probe with the complementary component when theaircraft is not in a nose-down attitude.

BRIEF DESCRIPTION OF THE DRAWINGS

Referring now to the drawings wherein like elements are numbered alikein the several FIGURES:

FIG. 1 depicts a rotary wing aircraft in an exemplary embodiment;

FIG. 2 is a perspective view of a rotary wing aircraft in an exemplaryembodiment;

FIG. 2A depicts a planform of a rotor blade in an exemplary embodiment;

FIG. 3 is a perspective view of a gear train for a rotary wing aircraftin an exemplary embodiment;

FIGS. 3A and 3B depict power distribution in the gear box in hover andcruise modes in exemplary embodiments;

FIG. 3C depicts plots of percentage of power versus airspeed for a mainrotor assembly and a propeller in an exemplary embodiment;

FIG. 3D depicts plots of percentage of lift versus airspeed for a mainrotor assembly and a propeller in an exemplary embodiment;

FIG. 4 is a perspective view of a gearbox and translational thrustsystem in an exemplary embodiment;

FIG. 5 is a perspective view of a rotor hub fairing in an exemplaryembodiment;

FIG. 6 depicts a flight control system in an exemplary embodiment;

FIG. 6A depicts a blade proximity detection system in an exemplaryembodiment;

FIG. 7 depicts a flight maneuver in an exemplary embodiment;

FIG. 8 depicts front, side and top views of an aircraft in an exemplaryembodiment;

FIG. 9 depicts an active vibration control (AVC) system in an exemplaryembodiment:

FIGS. 10 and 11 illustrate force vectors in exemplary hover states; and

FIG. 12 depicts a side view of a rotary wing aircraft including an airrefueling nozzle in an extended position in an exemplary embodiment.

DETAILED DESCRIPTION

FIG. 1 depicts an exemplary embodiment of a rotary wing, verticaltakeoff and land (VTOL) aircraft 10. The aircraft 10 includes anairframe 12 with an extending tail 14. A dual, counter rotating, coaxialmain rotor assembly 18 is located at the airframe 12 and rotates about amain rotor axis, A. In an exemplary embodiment, the airframe 12 includestwo seats for flight crew (e.g., pilot and co-pilot) and six seats forpassengers. The main rotor assembly 18 is driven by a power source, forexample, one or more engines 24 via a gearbox 26. The main rotorassembly 18 includes an upper rotor assembly 28 driven in a firstdirection (e.g., counter-clockwise) about the main rotor axis, A, and alower rotor assembly 32 driven in a second direction (e.g., clockwise)about the main rotor axis, A, opposite to the first direction (i.e.,counter rotating rotors). Each of the upper rotor assembly 28 and thelower rotor assembly 32 includes a plurality of rotor blades 36 securedto a rotor hub 38. In some embodiments, the aircraft 10 further includesa translational thrust system 40 located at the extending tail 14 toprovide translational thrust (forward or rearward) for aircraft 10.

Any number of blades 36 may be used with the rotor assembly 18. FIG. 2Adepicts a planform of a rotor blade 36 in an exemplary embodiment. Therotor assembly 18 includes a rotor hub fairing 37 generally locatedbetween and around the upper and lower rotor assemblies such that therotor hubs 38 are at least partially contained therein. The rotor hubfairing 37 provides drag reduction. Rotor blades 36 are connected to theupper and lower rotor hubs 38 in a hingeless manner, also referred to asa rigid rotor system. Although a particular aircraft configuration isillustrated in this non-limiting embodiment, other rotary-wing aircraftwill also benefit from embodiments of the invention. Although, the dualrotor system is depicted as coaxial, embodiments include dual rotoraircraft having non-coaxial rotors.

The translational thrust system 40 includes a propeller 42 connected toand driven by the engine 24 via the gearbox 26. The translational thrustsystem 40 may be mounted to the rear of the airframe 12 with atranslational thrust axis, T, oriented substantially horizontal andparallel to the aircraft longitudinal axis, L, to provide thrust forhigh-speed flight. The translational thrust axis, T, corresponds to theaxis of rotation of propeller 42. While shown in the context of apusher-prop configuration, it is understood that the propeller 42 couldalso be more conventional puller prop or could be variably facing so asto provide yaw control in addition to or instead of translationalthrust. It should be understood that any such system or othertranslational thrust systems may alternatively or additionally beutilized. Alternative translational thrust systems may include differentpropulsion forms, such as a jet engine.

Referring to FIG. 2, translational thrust system 40 includes a propeller42 and is positioned at a tail section 41 of the aircraft 10. Propeller42 includes a plurality of blades 47. In exemplary embodiments, thepitch of propeller blades 47 may be altered to change the direction ofthrust (e.g., forward or rearward). The tail section 41 includes activeelevators 43 and active rudders 45 as controllable surfaces.

Shown in FIG. 3 is a perspective view of portions of main rotor assembly18 and gearbox 26. The gearbox 26 includes an upper bull gear 44 rotatesabout the main rotor axis, A, and connected to the lower rotor assembly32 via a lower rotor shaft 46 extending along the main rotor axis, A. Alower bull gear 48 rotates about the main rotor axis, A, and isconnected to the upper rotor assembly 28 via an upper rotor shaft 50extending along the main rotor axis, A, and through an interior of thelower rotor shaft 46. Torque and rotational speed are provided to thegearbox 26 via input shaft 52 that transmits the torque and rotationalspeed from the engine(s) 24 to an input bevel gear 54 disposed at aninput bevel shaft 56 of the gearbox 26 via an input bevel pinion 104. Insome embodiments, the input bevel shaft 56 rotates about an input bevelshaft axis 58 parallel to the main rotor axis A. The propeller 42 isdriven by a propeller output shaft 106 driven by a propeller output gear62 disposed at a quill shaft 102, or an extension of input bevel shaft56. Transfer from the propeller output gear 62 is achieved viaconnection with a propeller output pinion 60 at the propeller outputshaft 106. To transfer torque from the input bevel shaft 56 to the lowerrotor assembly 32 and the upper rotor assembly 30, the gearbox 26includes a torque split gear reduction stage 64. The torque split gearreduction stage 64 splits torque from the input shaft 52 and applies thedivided torque to bull gears 44 and 48, respectively. While shown withthe propeller output shaft 106 driven by the propeller output gear 62,it is understood that such elements could be removed where the propeller42 is not used or is separately driven.

FIG. 3A illustrates power distribution through gearbox 26 to main rotorassembly 18 and propeller output shaft 106 during hover mode. In hover,power flows to torque split section to drive main rotor assembly 18. Thepropeller output shaft 106 spins at all times to drive features onpropeller box while propeller 42 is unclutched. During hover mode, themajority of power flows to the main rotor assembly 18.

FIG. 3B illustrates power distribution through gearbox 26 to main rotorassembly 18 and propeller output shaft 106 during cruise mode. In highspeed cruise, the majority of power flows to the propeller output shaft106 while the main rotor assembly 18 is operating near an autorotativestate.

FIG. 3C depicts a plot of percentage of power versus airspeed for themain rotor assembly 18 and the propeller 42. The power between the mainrotor assembly 18 and the propeller 42 is inversely proportional to airspeed, once the aircraft reaches a propeller engagement speed. Forexample, at low airspeeds (e.g. below 100 kts), power is 100% used bythe main rotor assembly 18. At the transition speed where the propeller42 engages, the propeller 42 begins to use aircraft power. As airspeedincreases, the main rotor assembly 18 power decreases and the propeller42 power increases.

FIG. 3D depicts plots of percentage of lift versus airspeed for the mainrotor assembly 18 and the propeller 42 in an exemplary embodiment. Whenaircraft 10 is flown in a nose up orientation, lift may be provided fromthe propeller 42. The lift supplied by the main rotor assembly 18 andthe propeller 42 is inversely proportional to air speed, once theaircraft reaches a propeller engagement speed. For example, at lowairspeeds (e.g. below 100 kts), lift is 100% provided by the main rotorassembly 18. At the transition speed where the propeller 42 engages, thepropeller 42 begins to provide lift. As airspeed increases, the mainrotor assembly 18 lift decreases and the propeller 42 lift increases.

Referring to FIG. 4, the main rotor assembly 18 is driven about the axisof rotation, A, through a main gearbox (MGB) 26 by a multi-enginepowerplant system 24, having two engine packages ENG1, ENG2 in theexample in FIG. 4. Although FIG. 4 depicts two engines 24, it isunderstood that aircraft 10 may use a single engine 24. The multi-enginepowerplant system 24 generates power available for flight operations andcouples such power to the main rotor assembly 18 and the translationalthrust system 40 through the MGB 26. The MGB 26 may be interposedbetween the powerplant system 24, the main rotor assembly 18 and thetranslational thrust system 40.

A portion of the drive system downstream of the MGB 26 includes acombined gearbox 90 (also referred to as a clutch). The combined gearbox90 selectively operates as a clutch and a brake for operation of thetranslational thrust system 40 with the MGB 26. The combined gearbox 90also operates to provide a rotor brake function for the main rotorassembly 18.

The combined gearbox 90 generally includes an input 92 and an output 94generally defined along an axis parallel to rotational axis, T. Theinput 92 is generally upstream of the combined gearbox 90 relative theMGB 26 and the output 94 is downstream of the combined gearbox 90 andupstream of the pusher propeller system 40 (FIG. 2). The combinedgearbox 90 may be categorized by the technique used to disengage-engage(e.g., clutch) or stop (e.g., brake) the load such as friction,electromagnetic, mechanical lockup, etc., and by the method used toactuate such as mechanical, electric, pneumatic, hydraulic,self-activating, etc. It should be understood that various combinedgearbox 90 systems may be utilized to include but not to be limited tomechanical, electrically, hydraulic and various combinations thereof.

Referring to FIG. 5, an exemplary rotor hub fairing 37 is shown. Rotorhub fairing 37 is illustrated having generally elliptical, incross-section, upper and lower hub fairings 111 and 112, and anairfoil-type shape (in horizontal cross-section) for the shaft fairing103. The airfoil shape of the shaft fairing 103 includes a leading edge114, and a trailing edge 115 aft of the upper and lower fairings 111,112. The airfoil shape of the shaft fairing 103 additionally includes achord (not shown) that connects the leading and trailing edges 114, 115of the airfoil. In one embodiment, the airfoil shape, including theupper surface 116 and the lower surface 117, is symmetrical about aplane extending along the length of the shaft fairing 103 and containingthe axis of rotation, A. As noted above, the upper and lower rotor hubs38 may be positioned, at least partially, in the upper and lowerfairings 111, 112.

Portions of the aircraft 10 are controlled by a flight control system120 illustrated in FIG. 6. In one embodiment, the flight control system120 is a fly-by-wire (FBW) control system. In a FBW control system thereis no direct mechanical coupling between a pilot's controls and movablecomponents of aircraft 10. Instead of using mechanical linkages, a FBWcontrol system includes a plurality of sensors 122 which can sense theposition of controlled elements and generate electrical signalsproportional to the sensed position. The sensors 122 may also be useddirectly and indirectly to provide a variety of aircraft state data to aflight control computer (FCC) 124. The FCC 124 may also receive inputs126 as control commands from various sources. For instance, the inputs126 can be pilot inputs, auto-pilot inputs, navigation system basedinputs, or any control inputs from one or more control loops executed bythe FCC 124 or other subsystems. In response to inputs from the sensors122 and inputs 126, the FCC 124 transmits signals to various subsystemsof the aircraft 10.

Flight control system 120 may include a rotor interface 128 configuredto receive commands from the FCC 124 and control one or more actuators,such as a mechanical-hydraulic or electric actuators, for the upperrotor assembly 28 and lower rotor assembly 32. In an embodiment, inputs126 including cyclic, collective, pitch rate, and throttle commands thatmay result in the rotor interface 128 driving the one or more actuatorsto adjust upper and lower swashplate assemblies (not depicted) for pitchcontrol of the upper rotor assembly 28 and lower rotor assembly 32.Alternatively, pitch control can be performed without a swashplateassemblies using individual blade control (IBC) in the upper rotorassembly 28 and lower rotor assembly 32. The rotor interface 128 canmanipulate the upper rotor assembly 28 and lower rotor assembly 32independently. This allows different collective and cyclic commands tobe provided to the upper rotor assembly 28 and lower rotor assembly 32.

Flight control system 120 may include a translational thrust interface130 configured to receive commands from the FCC 124 to control one ormore actuators, such as a mechanical-hydraulic or electric actuators,for the control of the translational thrust system 40. In an embodiment,inputs 126 may result in the translational thrust interface 130controlling speed of propeller 42, altering the pitch of propellerblades 47 (e.g., forward or rearward thrust), altering the direction ofrotation of propeller 42, controlling gearbox 90 to employ a clutch toengage or disengage the propeller 42, etc.

Flight control system 120 may include a tail fairing interface 132. Thetail fairing interface 132 is configured to receive commands from theFCC 124 to control one or more actuators, such as a mechanical-hydraulicor electric actuators, for the active elevator 43 and/or active rudders45 of FIG. 2. In an embodiment, inputs 126 include an elevator pitchrate command for the tail fairing interface 132 to drive the one or moreactuators for pitch control of the active elevators 43 of FIG. 2. In anembodiment, inputs 126 include a rudder command for the tail fairinginterface 132 to drive the one or more actuators for positional controlof the active rudders 45 of FIG. 2.

Flight control system 120 may include an engine interface 133. Theengine interface 133 is configured to receive commands from the FCC 124to control engine(s) 24. In an embodiment, inputs 126 include a throttlecommand from the pilot to adjust the RPM of engine(s) 24. FCC 124 mayalso send commands to engine interface 133 to control the engine(s) incertain predefined operating modes (e.g., quiet mode).

The FCC 124 includes a processing system 134 that applies models andcontrol laws to augment commands based on aircraft state data. Theprocessing system 134 includes processing circuitry 136, memory 138, andan input/output (I/O) interface 140. The processing circuitry 136 may beany type or combination of computer processors, such as amicroprocessor, microcontroller, digital signal processor, applicationspecific integrated circuit, programmable logic device, and/or fieldprogrammable gate array, and is generally referred to as centralprocessing unit (CPU) 136. The memory 138 can include volatile andnon-volatile memory, such as random access memory (RAM), read onlymemory (ROM), or other electronic, optical, magnetic, or any othercomputer readable storage medium onto which data and control logic asdescribed herein are stored. Therefore, the memory 138 is a tangiblestorage medium where instructions executable by the processing circuitry136 are embodied in a non-transitory form. The I/O interface 140 caninclude a variety of input interfaces, output interfaces, communicationinterfaces and support circuitry to acquire data from the sensors 122,inputs 126, and other sources (not depicted) and communicate with therotor interface 128, the translation thrust interface 130, tail faringinterface 132, engine interface 133, and other subsystems (notdepicted).

In exemplary embodiments, the rotor interface 128, under control of theFCC 124, can control the upper rotor assembly 28 and lower rotorassembly 32 to pitch in different magnitudes and/or different directionsat the same time. This includes differential collective, where the upperrotor assembly 28 has a collective pitch different than the collectivepitch of the lower rotor assembly 32, in magnitude and/or direction.Differential pitch control also includes differential cyclic pitchcontrol, where the upper rotor assembly 28 has a cyclic pitch differentthan the cyclic pitch of the lower rotor assembly 32, in magnitude, axisof orientation (e.g., longitudinal or lateral) and/or direction. Thedifferential collective and the differential cyclic pitch control may beaccomplished using independently controlled swashplates in the upperrotor assembly 28 and lower rotor assembly 32. Alternatively,differential collective and the differential cyclic pitch control may beaccomplished using individual blade control in the upper rotor assembly28 and lower rotor assembly 32.

The ability to independently control the pitch of the upper rotorassembly 28 and lower rotor assembly 32 allows the lower rotor assembly32 to be adjusted due to its position beneath the upper rotor assembly28. The lower rotor assembly 32 is located in the downwash of the upperrotor assembly 28. To accommodate for this, the lower rotor assembly 32may have a collective pitch that differs from the collective pitch ofthe upper rotor assembly 28.

In the case of traditional helicopters, as the forward velocity of theaircraft increases, the velocity of the retreating blade relative to theairflow decreases. This causes a stall region to arise at the root ofthe retreating blade and expand towards to distal end of the blade asspeed increases. As this stall region increases, the overall lift vectorof the aircraft shifts from the center of the aircraft towards theadvancing blade which is providing the majority of lift for theaircraft. This imbalance of lift creates an unstable rolling moment onthe aircraft which is stabilized by a combination of reducing forwardflight and blade flapping, which reduces overall aircraft lift. With adual rotor aircraft, such as aircraft 10, the counter rotating rotorheads balance out the torque generated by each rotor head and alsobalances the lift generated by each advancing blade without the need forblade flapping or reducing the speed of the aircraft. This is madepossible by the rigid rotor system. With two rigid rotors, the rollmoments cancel at the main rotor shaft.

The use of upper rotor assembly 28 and lower rotor assembly 32 allowsthe pre-cone angle to be set on each individual rotor to reduce bendingstress on the blades. In a hinged rotor design, the hinges willnaturally go to an angle to reduce bending stress. On a rigid rotoraircraft, such as aircraft 10, there is no hinge, so the pre-cone angleis set to avoid the extra stress attributed to the bending moment. Auseful pre-cone angle is one where the centrifugal force of the bladepulling out matches the lift of the blade up. Due to the independentnature of the upper rotor assembly 28 and lower rotor assembly 32,differential pre-cone is used in aircraft 10. Differential pre-conerefers to the fact that the upper rotor assembly 28 and lower rotorassembly 32 have different pre-cone angles. The different pre-coneangles for the upper rotor assembly 28 and lower rotor assembly 32 helpmaintain tip clearance. In an exemplary embodiment, the pre-angle on theupper rotor assembly 28 is about 3 degrees and the pre-cone angle on thelower rotor assembly 32 is about 2 degrees.

Aircraft 10 is operational in a variety of modes, including take-off,cruise, landing, etc. Cruise mode refers to generally horizontal flight.During cruise, aircraft 10 can reach speeds of above about 200 knots,with speed reaching up to about 250 knots. During cruise mode, the mainrotor assembly 18 provides the majority of lift for the aircraft. Inexemplary embodiments and flight modes, the main rotor assembly 18provides greater than about 85% of the lift during cruise mode.

Aircraft 10 may assume various acoustic modes, depending on the flightstate. FCC 124 may control RPM of engines 24, RPM of propeller 42, andclutch 90 to engage or disengage the propeller 42 to assume differentnoise levels. For example, at take-off noise may not be a concern, andthere would be no changes in aircraft operation to adjust the noiselevel. As the aircraft approaches a target, it may be desirable todisengage the propeller 42 using clutch 90 and/or reduce RPM of engines24 to reduce the noise produced by aircraft 10. The propeller 42 may bedisengaged at various other flight states (e.g., high speed) to reducenoise. The RPM of the main rotor assembly 18 and RPM of propeller 42 maybe independently controlled (e.g., through clutch 90 or FCC 124). Thisallows a variety of flight states to be achieved.

The pilot may enter separate commands to reduce aircraft noise, forexample, disengaging the propeller 42, reducing engine RPM, andincreasing collective pitch as separate inputs. Alternatively, the pilotmay select a reduced noise mode (e.g., quiet mode) through single input,and the FCC 124 controls the various aircraft interfaces to achieve thedesired mode. For example, the pilot may select a reduced noise mode atinput 126, and the FCC automatically disengages the propeller 42,reduces the engine 24 RPM and/or increases collective pitch withoutfurther demand on the pilot.

The use of the translational thrust system 40 allows the aircraft 10 tomove forward or rearward (depending on the pitch of the propellerblades) independent of the pitch attitude of the aircraft. Cyclic isused to adjust the pitch attitude (nose up, nose down or level) of theaircraft while the translational thrust system 40 provides forward andrearward thrust.

The motor rotor assembly 18 system and the translational thrust system40 are connected through the main gear box 26. A gear ratio of main gearbox 26 is selected so as to keep propeller 42 at a high efficiency andsuitable noise level during cruise mode. The gear ratio of main gear box26 dictates the ratio of the rotor speed of main rotor assembly 18 topropeller speed of propeller 42.

Embodiments of aircraft 10 provide the pilot with increased situationalawareness by allowing the aircraft attitude (e.g., the angle oflongitudinal axis, L, relative to horizontal) to be adjusted by cyclicpitch of the main rotor assembly 18 and the forward and rearward thrustto be controlled by the translational thrust system 40. This allows avariety of flight modes to be achieved, which allows the pilot to bemore aware of their surroundings. Aircraft 10 can take off at ahorizontal attitude (e.g., axis L is horizontal), which also may bereferred to as vertical take-off. Aircraft 10 may also fly forward orcruise with the nose angled upwards, nose angled downwards or level.Aircraft 10 can hover with the nose angled upwards or downwards orlevel. FIGS. 10 and 11 illustrate force vectors from the main rotorassembly and propeller for hover nose up and hover nose down,respectively. Aircraft 10 can also land substantially parallel to anon-horizontal or sloped surface by adjusting the pitch attitude of theaircraft using cyclic pitch of the main rotor assembly 18. The use ofmain rotor assembly 18 for aircraft attitude and the translationalthrust system 40 for thrust allows aircraft 10 to assume a variety oftrim states.

Embodiments provide independent control of the active elevators 43and/or active rudders 45 as controllable surfaces in the tail section41. The elevator surfaces 43 may be controlled independently by the FCC124 through the tail faring interface 132. The rudder surfaces 45 may becontrolled independently by the FCC 124 through the tail faringinterface 132.

The configuration of aircraft 10 and the controlled afforded by FCC 124allows aircraft 10 to provide a high bank angle capability at highspeeds. For example, in an exemplary embodiment, aircraft 10 canmaintain a bank angle of about 60 degrees at about 210 knots.

Aircraft 10 may make use of longitudinal lift offset in trim tocompensate for rotor-on-rotor aerodynamic interaction between the upperrotor assembly 28 and lower rotor assembly 32. Aircraft 10 may adjustdifferential longitudinal cyclic as a function of operational states ofthe aircraft (e.g., take-off, cruise, land, etc.). Differentiallongitudinal cyclic refers to upper rotor assembly 28 and lower rotorassembly 32 having different cyclic pitch along the longitudinal axis ofthe aircraft. Differential longitudinal cyclic may also be used togenerate yaw moments. Lift offset may be used to control aircraft, wherelateral lift offset adjusts roll and longitudinal lift offset adjustspitch.

FCC 124 may control RPM of engine(s) 24, RPM of propeller 42, and clutch90 to engage or disengage the propeller 42 to assume different noiselevels. For example, at take-off noise may not be a concern, and therewould be no changes in aircraft operation to adjust the noise level. Asthe aircraft approaches a target, it may be desirable to disengage thepropeller 42 using clutch 90 and/or reduce RPM of engines 24 to reducethe noise produced by aircraft 10. The propeller 42 may be disengaged atvarious other flight states (e.g., high speed) to reduce noise. The RPMof the main rotor assembly 18 and RPM of propeller 42 may beindependently controlled (e.g., through clutch 90).

The pilot may enter separate commands to reduce aircraft noise, forexample, disengaging the propeller 42 and reducing engine RPM asseparate inputs. Alternatively, the pilot may select a reduced noisemode (e.g., quiet mode) through single input, and the FCC 124 controlsthe various aircraft interfaces to achieve the desired mode. Forexample, the pilot may select a reduced noise mode at input 126, and theFCC automatically disengages the propeller 42 and/or reduces the engine24 RPM without further demand on the pilot.

Aircraft 10 provides the ability to approach a target and reverse thrustwhile maintaining an attitude directed at the target. FIG. 7 depictsaircraft 10 approaching a target 200. In a first state, 202, theaircraft 10 alters the pitch of blades 47 in propeller 42 to providereverse thrust to bring the aircraft to a quick stop. At state 204, themain rotor assembly 18 and propeller 42 are controlled to pitch aircraft10 towards target 200. At state 206, the propeller 42 is used to providereverse thrust to move away from target 200 and climb, while stillmaintaining an attitude with the nose of aircraft 10 facing target 200.

The use of a dual rotor system and translational thrust allows aircraft10 to eliminate the need for a variable angle between the main axis ofrotation of the rotor system (e.g., axis A in FIG. 1) and aircraftlongitudinal axis L. In conventional helicopters, the angle between themain axis of rotation of the rotor system and the aircraft longitudinalaxis L varies. This is due to the fact that conventional helicopterslack a translational thrust system 40 for use during cruise mode, orforward flight. In a conventional helicopter, forward flight is providedthrough cyclic pitch, which causes the aircraft to point nose down. Asthis nose down orientation is undesirable beyond a certain angle, theangle between the main axis of rotation of the rotor system and theaircraft longitudinal axis L is adjusted to bring the nose upwards,while still in forward flight.

By contrast, aircraft 10, with translational thrust system 40, does notneed to adjust the angle between the main axis of rotation of the rotorsystem (e.g., axis A in FIG. 1) and aircraft longitudinal axis L. Theangle between the main axis of rotation of the rotor system (e.g., axisA in FIG. 1) and aircraft longitudinal axis L for aircraft 10 remainsfixed during all flight modes, including take-off, cruise, landing, etc.In other embodiments, a pitch attitude of the translational thrustsystem 40 and/or the rotor system may be used to achieve maximumacceleration.

As shown in FIG. 1, the rotor assembly 18 includes a rotor hub fairing37 generally located between and around the upper and lower rotorassemblies such that the rotor hubs 38 are at least partially containedtherein. The rotor hub fairing 37 provides drag reduction. Referring toFIG. 5, an exemplary rotor hub fairing 37 is shown. Rotor hub fairing 37is illustrated having generally elliptical, in cross-section, upper andlower hub fairings 111 and 112, and an airfoil-type shape (in horizontalcross-section) for the shaft fairing 103. The airfoil shape of the shaftfairing 103 includes a leading edge 114, and a trailing edge 115 aft ofthe upper and lower fairings 111, 112. The airfoil shape of the shaftfairing 103 additionally includes a chord (not shown) that connects theleading and trailing edges 114, 115 of the airfoil. In one embodiment,the airfoil shape, including the upper surface 116 and the lower surface117, is symmetrical about a plane extending along the length of theshaft fairing 103 and containing the axis of rotation, A. As notedabove, the upper and lower rotor hubs 38 may be positioned, at leastpartially, in the upper and lower fairings 111, 112.

The rotor hub fairing 37 is a sealed fairing, meaning there are few orno passages for air to travel through the interior of the rotor hubfairing 37. In conventional designs, control devices such as pushrods,are exposed near the rotor hubs. The surfaces of these componentsincrease drag on the rotor assembly. The air gaps between various rotorstructures (e.g., pushrods and main rotor shaft) also form areas ofdrag. The sealed rotor hub fairing 37 eliminates air pathways throughthe rotor hub structure, and eliminates drag associated with such airpaths.

Another feature to reduce drag on the rotor hub is positioning controlrods, such as push rods for rotor control, internal to the main rotorshaft. Referring to FIG. 3, pushrods for swashplates in the upper rotorassembly 28 and lower rotor assembly 32 are located internal to thelower rotor shaft 46 and upper rotor shaft 50. This prevents thepushrods from being exposed and increasing drag on the rotor hub. Theuse of a rigid rotor system aids in sealing the rotor hub faring 37.

In an exemplary embodiment, the distance between the hub of the upperrotor assembly 28 and the hub of the lower rotor assembly 32 ranges fromabout 2 feet to about 2.5 feet. In another exemplary embodiment, thedistance between the hub of the upper rotor assembly 28 and the hub ofthe lower rotor assembly 32 ranges from about 2.1 feet to about 2.4feet. In another exemplary embodiment, the distance between the hub ofthe upper rotor assembly 28 and the hub of the lower rotor assembly 32is about 2.29 feet (0.7 meters).

Aircraft 10 may employ an active vibration control (AVC) system toreduce vibration in the airframe 12. The use of a dual rotor, rigidrotor system tends to produce significant vibration in the airframe 12and its systems. FIG. 9 depicts an AVC system in an exemplaryembodiment. An AVC controller 300 executes an AVC control process toreduce vibration in aircraft 10. AVC controller 300 may be implementedas part of flight control system 120, executed by FCC 124, or may be aseparate controller. One or more sensors 302 are located in aircraft 10to detect vibration. Sensors may be located in a wide variety ofpositions, including airframe 12, gearbox 26, tail section 14, on mainrotor assembly 18, cockpit, etc. It is understood that these locationsare exemplary, and the AVC sensors 302 may be located in any position.AVC actuators 304 generate a force to dampen vibration in aircraft 10,as known in the art. AVC actuators 304 may be located in any position inthe aircraft.

In operation, AVC controller 300 receives vibration signals from the AVCsensors 302. AVC controller 300 provides control signals to the AVCactuators 304 to generate forces to reduce the vibration sensed by theAVC sensors 302. Control signals to the AVC actuators 304 may vary inmagnitude and frequency to cancel vibrations in aircraft 10. In anexemplary embodiment, AVC controller 300 operates in a feedback mode,where the control signals to AVC actuators 304 are adjusted in responseto measured vibration from AVC sensors 302. In an alternate embodiment,AVC controller 300 does not actively measure vibration through AVCsensors 302. Rather, the AVC controller 300 obtains the rotor speed(e.g., through an RPM signal) and applies a control signal to the AVCactuators 304, in an open loop control mode.

The use of independently controlled upper rotor assembly 28 and thelower rotor assembly 32, along with other control surfaces, provides theability to control yaw using a variety of elements. For example, below afirst speed, (e.g., 40 knots), the FCC 124 uses differential collectivepitch for yaw control. Above the first speed but below a second speed(e.g., 80 knots), a mix of differential collective and differentialcyclic may be used to control yaw. The differential cyclic may beapplied along the longitudinal and/or lateral axes of the aircraft.Further, wind direction may be measured by a sensor 122 and used toadjust the differential cyclic about the longitudinal and/or lateralaxes. Above the second speed (e.g., 80 knots), the active rudders 45 areused as controllable surfaces to control yaw. The FCC 124 providescommands to the tail fairing interface 132 to control the rudders 45 toadjust yaw.

The use of active elevator 43, with independent control of a leftelevator section and a right elevator section, provides for improvedstability control. Flight control system 120 performs mixing ofcollective pitch of main rotor assembly 18 and an angle of elevator 43to provide stability augmentation.

Embodiments may use wireless techniques to provide tip clearancemeasurements. FIG. 6A depicts a blade proximity monitoring system in anexemplary embodiment. At least one upper rotor blade and at least onelower rotor blade is equipped with at least one antenna 502. Antennas502 may be electric field antennas or magnetic field antennas. Antennas502 may be implemented using compact ferrite core or small diametermagnet wire in the form of coils around the blade spar or embedded inthe plane of the blade skin. The antennas 502 interact through the nearfield effect.

An oscillator 504 sends an excitation signal (e.g., 40 KHz) to a firstantenna 502L. It is understood that the excitation signal may be sent toa plurality of antennas in different blades, including multiple antennasin the same blade. As the blades cross, a second antenna, 502U, receivesa signal emitted by the first antenna 502 L. An output level monitor 506measures the magnitude of the excitation signal.

A blade proximity monitor 508 (e.g., a processor implemented controller)is mounted in the rotating system, e.g., in a rotor hub. This eliminatesnoise that may be introduced through a conventional slip ring used toconvey signals from a rotating system to a stationary system. The bladeproximity monitor 508 receives an output signal from the second antenna502U and the magnitude of the excitation signal from the output levelmonitor 506. Output signal from the second antenna 502U may beamplified. The blade proximity monitor 508 also receives a RPM signal ofthe main rotor assembly 18 from a contactor 510. Based on the magnitudeof the excitation signal applied to the first antenna 502L and themagnitude of the output signal from the second antenna 502U, bladeproximity monitor 508 can detect the distance between the first antenna502L and the second antenna 502U. This provides an indication of thedistance between the rotor blades. The larger the magnitude of theoutput signal from second antenna 502U, the closer the blades.

The blade proximity monitor 508 may output the measured distance betweenthe blades to a rotor track and balance unit 512. The blade proximitymonitor 508 may output the measured distance between the blades toinstrument system 514 and to a pilot display 516. If the measureddistance goes below a threshold, then an alert may be generated to thepilot that the blades of the upper rotor assembly 32 and the lower rotorassembly 28 are too close to each other.

The use of a dual rotor, main rotor assembly 18 allows improvements incontrol of main rotor assembly 18. Flight control system 120 may applydifferent control envelopes to the upper rotor assembly 28 and the lowerrotor assembly 32. Flight control system 120 may impose differentcontrol ranges the upper rotor assembly 28 and the lower rotor assembly32 including control elements such as prioritization, gain vs.differential, collective versus cyclic, etc. The upper rotor assembly 28and the lower rotor assembly 32 may be independently controlled throughthe use of separate upper and lower swashplates. Alternatively, theupper rotor assembly 28 and the lower rotor assembly 32 may beindependently controller using individual blade control (IBC)techniques.

Aircraft 10 employs a fly-by-wire (FBW) control system to reduce pilotwork load. In an exemplary embodiment, FCC 124 determines the aircraftairspeed based on one or more sensors 122. The FCC 124 then adjusts thecollective pitch of the upper rotor assembly 28 and/or the lower rotorassembly 32 in response to the airspeed. FCC 124 may use a look up tablethat indexes airspeed to collective pitch. Alternatively, FCC 124 mayuse an algorithm to compute the collective pitch based on airspeed. Asnoted above, the collective pitch of upper rotor assembly 28 and thelower rotor assembly 32 may be the same or different.

Another feature to reduce pilot workload includes automaticallyadjusting the RPM and/or pitch of propeller 42 in response to a velocityor acceleration command from the pilot. Conventional systems wouldrequire the pilot to adjust propeller RPM and/or pitch throughindividual inputs. The flight control system 120 allows the pilot toenter a desired velocity or an acceleration, and the FCC 124 generatesthe proper commands to the translational thrust interface 130 toestablish an RPM and/or pitch to meet the desired velocity oracceleration.

In exemplary embodiments, the flight control system 120 controls themain rotor assembly 18 to prevent the tips of rotor blades 36 fromexceeding a threshold speed. In exemplary embodiments, the thresholdspeed may be 0.9 Mach 1. This threshold would prevent the rotor bladetips from exceeding the speed of sound. The threshold speed may vary,and may be set to limit drag on the rotor blades to below a certainlevel. In one embodiment, the FCC 124 determines air temperature fromsensors 122. FCC 124 may also determine prevailing wind speed anddirection from sensors 122. The FCC 124 then computes the thresholdspeed based on the speed of sound (e.g., Mach 1) at the sensed airtemperature. The FCC 124 may set the threshold to 0.9 Mach 1, forexample. FCC 124 then controls RPM of the main rotor assembly 18 toprevent the rotor blade tips from exceeding the threshold. In anexemplary embodiment, the FCC maintain 85% of the nominal rotor RPM. FCC124 may take into account prevailing wind direction and speed incontrolling the RPM of the main rotor assembly 18. The 0.9 Mach 1threshold is only one example, and other speed thresholds may beemployed to achieve desired results (e.g., reduce drag).

In exemplary embodiments, active elevator 43 is configured andcontrolled to improve stability, compensating for forces such aspropeller torque and/or rotor downwash. Elevator 43 includes a leftelevator and a right elevator on opposite sides of the axis of rotationof the propeller 42. The left elevator and right elevator may beindependently controlled to assume different positions. The tail fairinginterface 132 is configured to receive commands from the FCC 124 tocontrol one or more actuators, such as a mechanical-hydraulic orelectric actuators, to position the left elevator and right elevatorindependently. This independent control of the left elevator and rightelevator aids in compensating propeller torque and/or rotor downwash.

The left elevator and right elevator may also have different physicalconfigurations to compensate for compensating propeller torque and/orrotor downwash. The left elevator and right elevator may be offsetrelative to each other along the longitudinal and/or lateral axes ofaircraft 10. Further, the left elevator and right elevator may havedifferent geometries where one of the left elevator and right elevatoris larger than then other along the longitudinal and/or lateral axes ofaircraft 10. The left elevator and right elevator may have differingaerodynamic surfaces (e.g., airfoils) as well.

The cockpit of aircraft includes a single, semi-active, collective input(e.g., stick) positioned between the two pilot seats.

Exemplary embodiments of aircraft 10 provide a much smaller footprintthan existing aircraft. This makes aircraft 10 well suited for missionsin confined terrain, urban settings, and shipboard operations. FIG. 8presents front, side and top views of an exemplary aircraft. One featurecontributing to the reduced footprint is the location of the main rotorshaft relative to the airframe 12. As shown in FIG. 1, the axis ofrotation A, of the main rotor assembly 18, intersects longitudinal axis,L, along a span of axis L, extending from the nose of the aircraft tothe tip of the hub of propeller 42. In an exemplary embodiment, the axisof rotation A is located at about a 44% station (STA) of the fuselage orairframe 12.

In an exemplary embodiment, there is about 5.2 inches from the mainrotor pylon to the blade hub centerline. In an exemplary embodiment,there is about 0.7 inch hub clearance to the main rotor pylon. In anexemplary embodiment, the rotor blades 36 extend beyond the nose of theaircraft by about 13 inches (0.33 meters). In an exemplary embodiment,rotor blades 36 extend beyond the nose of the aircraft by about 6.9% ofthe blade span, which may be about 188.″

In one embodiment, the aircraft 10 may be configured to perform arefueling operation “in-flight.” Referring now to FIG. 12, an aerialrefueling probe 70 may be mounted to a portion of the airframe 12 suchthat a distal or free end 72 of the probe 70 is arranged generallyadjacent a front end 13, for example the nose, of the airframe 12. Theprobe 70 is fluidly coupled to one or more fuel tanks (not shown) of theaircraft 10 and includes a generally hollow conduit 74 and a tip 76connectable, such as via insertion, to a complementary component forfuel transfer.

The probe 70 is configured to move or transform, such as by pivoting orfolding in a non-telescoping manner for example, between a retractedposition and an extended position (as shown in the FIG.). During normalflight, the refueling probe 70 is in the retracted position, where it ishoused within a cavity 15 formed in the airframe 12. By stowing theprobe 70 during normal flight, the retracted probe 70 has a negligibleimpact on drag of the aircraft 10. In the retracted position, the probe70 may be arranged generally parallel to the aircraft longitudinal axisL. In response to an input generated by a pilot, the probe 70 pops outof the cavity 15, thereby transforming to the extended position. In theextended position, the probe 70 may be arranged at an angle to thelongitudinal axis L of the aircraft 10.

To perform an “in-flight” fueling operation, the pilot moves the fuelingprobe 70 into an extended position and approaches the complementarycomponent, such as a drogue connected to an end of a hose trailing froma fuel tanker (not shown). The pilot must maneuver the aircraft 10 toinsert the tip 76 of the probe 70 into the drogue. Only once a completeconnection is made can fuel flow freely through the fuel flow passageformed between the tanker to the aircraft 10.

The use of a rigid rotor system, along with the rotor shaft position(e.g., axis A) allows for much easier air-to-air refueling. Morespecifically, the use of a rigid rotor system eliminates the need forthe aircraft 10 to have a nose-down attitude when connecting to a fueltanker. As a result, the connection between the aircraft and the fueltanker may be formed at faster speeds, thereby increasing the range offixed wing aircrafts that may be used as tankers. The stiff rotor blades36 also ease air-to-air refueling by reducing blade flapping, which mayresult in a blade contacting a tanker fuel line during the refuelingprocess. In addition, because the portion of the rotor blades 36 thatextends forward of the airframe 12 is minimal, the length and thereforethe complexity of the refueling probe 70 are significantly reduced.

The smaller probe 70 not only has a decreased weight, but also has areduced impact on drag and other performance characteristics of theaircraft 10 in both the stowed, retracted position and the extendedposition. When extended, the probe 70 has a drag approximately ⅓ of thatof conventional helicopter probes. Further, the vertical drag on theextended probe 70 is not affected by rotor downwash when the aircraft 10is hovering.

In an exemplary embodiment, the distance between the hub of the upperrotor assembly 28 and the hub of the lower rotor assembly 32 ranges fromabout 2 feet to about 2.5 feet. In another exemplary embodiment, thedistance between the hub of the upper rotor assembly 28 and the hub ofthe lower rotor assembly 32 ranges from about 2.1 feet to about 2.4feet. In another exemplary embodiment, the distance between the hub ofthe upper rotor assembly 28 and the hub of the lower rotor assembly 32is about 2.29 feet. In another exemplary embodiment, the distancebetween a midpoint of a blade in the upper rotor assembly 28 and amidpoint of a blade in the lower rotor assembly 32 is about 29.0 inches.In another exemplary embodiment, the distance between a tip of a bladein the upper rotor assembly 28 and a tip of a blade in the lower rotorassembly 32 is about 31.0 inches. In another exemplary embodiment, thedistance between the hub of the upper rotor assembly 28 and the hub ofthe lower rotor assembly 32 is about 14% of the blade span, which may beabout 188 inches.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the invention.While the description of the present invention has been presented forpurposes of illustration and description, it is not intended to beexhaustive or limited to the invention in the form disclosed. Manymodifications, variations, alterations, substitutions, or equivalentarrangement not hereto described will be apparent to those of ordinaryskill in the art without departing from the scope and spirit of theinvention. Additionally, while the various embodiment of the inventionhave been described, it is to be understood that aspects of theinvention may include only some of the described embodiments.Accordingly, the invention is not to be seen as limited by the foregoingdescription.

1. An aircraft comprising: an airframe; an extending tail; a counterrotating, coaxial main rotor assembly including an upper rotor assemblyand a lower rotor assembly, the upper rotor assembly and the lower rotorassembly including a plurality of blades; and a translational thrustsystem positioned at the extending tail, the translational thrust systemproviding translational thrust to the airframe; wherein the plurality ofblades extend beyond a nose of the aircraft by about 13 inches.
 2. Theaircraft according to claim 1, further comprising a refueling probemounted to the airframe and fluidly connected to a fuel tank of theaircraft, the refueling probe being configured to form a connection witha complementary component to form a fuel flow passage there through. 3.The aircraft according to claim 2, wherein a first end of the refuelingprobe does not extend beyond the plurality of rotor blades.
 4. Theaircraft according to claim 2, wherein the refueling probe is configuredto move between a retracted position and an extended position, andwherein such movement between the retracted position and the extendedposition does not include telescoping.
 5. The aircraft according toclaim 2, wherein in the retracted position, the refueling probe ishoused within a cavity formed in the airframe near the nose.
 6. Theaircraft according to claim 2, wherein in the extended position, therefueling probe has a reduced impact on drag of the aircraft.
 7. Theaircraft according to claim 2, wherein the aircraft is configured toconnect the refueling probe with the complementary component when theaircraft is not in a nose-down attitude.